The present invention relates generally to gas turbine engines, and, more specifically, to augmented engines.
Gas turbine engines are available in a variety of configurations, sizes, and output thrust. The typical turbofan engine includes in serial flow communication a fan, multistage compressor, combustor, high pressure turbine (HPT), and low pressure turbine (LPT). The HPT is joined to the compressor by one shaft, and the LPT is joined to the fan by another shaft.
In operation, air enters the engine and flows past the fan and into the compressor wherein it is pressurized and discharged to the combustor. Fuel is added to the compressed air in the combustor for generating hot combustion gases which are discharged through the turbine stages.
The HPT includes rotor blades which extract energy from the combustion gases for rotating the supporting disk joined to the shaft which powers the compressor. The combustion gases then flow to the LPT which includes additional turbine blades which extract additional energy from the gases for rotating the supporting rotor disk which powers the shaft to drive the upstream fan.
An annular bypass duct typically surrounds the core engine for bypassing a portion of the air pressurized by the fan for producing propulsion thrust independently from the core exhaust gases discharged from the core engine through the LPT.
The engine has a surrounding nacelle which may be relatively short for defining a separate fan exhaust nozzle located upstream from the exhaust nozzle for the core engine. The nacelle may also be long extending downstream to a common engine outlet through which both the fan bypass air and the core exhaust gases are discharged collectively for producing the propulsion thrust.
Gas turbine engines are highly sophisticated and complex and include many individual parts which must be separately manufactured and assembled in the final engine. The number of parts required in a particular engine directly affects the complexity thereof and the associated manufacturing cost.
Cost is always a primary consideration in manufacturing gas turbine engines. And, cost is particularly significant in manufacturing gas turbine engines for expendable applications, or for a limited number of uses. For example, the typical military cruise missile includes a small turbofan engine which is specifically configured for the one-time use of the missile which, of course, is destroyed upon reaching its target.
Furthermore, small, pilotless drones or remotely piloted vehicles may also benefit from the use of small turbofan engines for a limited number of flight missions.
The small turbofan engines used in these limited-mission applications are specifically configured for subsonic operation at less than the speed of sound. However, for applications requiring supersonic operation at greater than the speed of sound, the engines require additional propulsion thrust.
Such additional thrust is typically provided by introducing an afterburner or augmentor at the discharge end of the turbofan engine for burning additional fuel and generating additional propulsion thrust when desired. The typical afterburner is a complex assembly of many parts including flameholders, fuel injectors, a combustion liner, and a variable area exhaust nozzle.
The exhaust nozzle is particularly complex since it must be configured for operation in a dry mode for normal operation of the basic turbofan engine, as well as during a wet or reheat mode of operation when additional fuel is burned in the afterburner.
The variable area nozzle must have a specific axial profile which converges to a throat of minimum area and then diverges to the outlet of the nozzle in a commonly configured converging-diverging (C-D) nozzle.
The throat of the nozzle provides a minimum discharge flow area for dry operation of the engine. When the afterburner is activated additional fuel is consumed, which in turn requires a larger discharge throat area, as well as a suitable C-D nozzle profile for achieving supersonic exhaust velocity from the nozzle. The larger exhaust area during wet operation is required to prevent excessive back pressure on the core engine and, in particular the compressor thereof, to avoid undesirable compressor stall.
The C-D variable area nozzle is therefore substantially complex and requires individual exhaust flaps joined together to create the articulated C-D nozzle profile which may be varied between the dry and wet modes of operation. Suitable actuators and seals are also required to ensure proper performance of the nozzle and engine over a suitable life span.
In view of the complexity of the typical C-D nozzle, the introduction thereof in the relatively small expendable or limited use aircraft applications described above is not practical, if even possible.
Accordingly, it is desired to provide an expendable or limited use gas turbine engine with a simplified afterburner for achieving supersonic flight operation.